1. Field of the Invention
This invention relates to coolable walls such as film cooled combustor liners for use in gas turbine engines, and more particularly, to aircraft gas turbine engine afterburner liners.
The invention described herein was made in the performance of work under U.S. NAVY contract N00019-80-C-0017 and the U.S. Government has rights therein.
2. Description of Related Art
Afterburning aircraft gas turbine engines, such as turbofan engines used in fighter aircraft, burn fuel in the exhaust section of the engine producing a hot gas usually at an intensely high temperature such as 3,000.degree. F. or even higher. To prevent this intense heat from damaging the afterburner and surrounding parts of the engine, before the hot gas exits to an exhaust nozzle, an afterburner heat shield or combustor liner is provided.
Conventional afterburner combustion liners incorporate convection and film cooling that are provided by circumferentially disposed rows of film cooling slots such as those depicted in U.S. Pat. No. 4,566,280 by Burr and U.S. Pat. No. 4,733,538 by Vdoviak et al. These liners are typified by complex structures that have non-uniform liner thicknesses that give rise to thermal gradients that in turn cause low cycle fatigue in the liner and therefore shorten their potential life expectancy and reduce their durability. The complex shapes and fabrication required to produce these liners adversely affects their cost and weight.
A more detailed discussion of the related art may be found in related U.S. Pat. No. 5,181,379 entitled "GAS TURBINE ENGINE MULTI-HOLE FILM COOLED COMBUSTOR LINER AND METHOD OF MANUFACTURE", invented by Wakeman et al., which issued Jan. 26, 1993, assigned to the same assignee and U.S. Pat. No. 5,233,828 entitled "COMBUSTOR LINER WITH CIRCUMFERENTIALLY ANGLED FILM COOLING HOLES", invented by Napoli, which issued Aug. 10, 1993, assigned to the same assignee, and both are incorporated herein by reference.
Engine designers have long sought to incorporate low weight single wall afterburner liners capable of withstanding the temperatures and pressure differentials found in afterburners. To that end the invention described in the Wakeman reference provides a single wall, preferably sheet metal, annular combustor liner having multi-hole film cooling holes which are disposed through the wall of the liner at sharp downstream angles. The multi-hole film cooling holes are spaced closely together to form at least one continuous pattern designed to provide film cooling over the length of the liner. The Wakeman reference discloses multi-hole film cooling holes which have a diameter of about 20 mils with a nominal tolerance of about .+-.2 mils, are spaced closely together about 61/2to 71/2hole diameters apart, have a downstream angle of 20 degrees with a nominal tolerance of about .+-.1 degree. Axially adjacent holes are circumferential offset by half the angle (or distance) between circumferentially adjacent holes to 10 further enhance the evenness of the cooling film injection points. The Wakeman reference further discloses an embodiment wherein the liner may be corrugated so as to form a wavy wall which is designed to prevent buckling and is particularly useful for aircraft gas turbine engine afterburners.
Afterburner cooling uses a significant percentage of cooling air which is usually taken from the fan section of the engine, This is costly air in terms of fuel and power consumption and therefore aircraft engine designers are always seeking means for reducing the amount of cooling air required to cool afterburner liners.